Compressor stator with leading edge fillet

ABSTRACT

A compressor of a gas turbine engine includes a rotor and a stator located downstream of the rotor. The stator has a plurality of vanes each with an airfoil extending span-wise between a root proximate an inner hub of the stator and a tip. A fillet is disposed at the leading edge of the root of the airfoil, and extends between a pressure side surface of the airfoil and the inner hub.

TECHNICAL FIELD

The application relates generally to stators for gas turbine engines,and more particularly, to compressor stators.

BACKGROUND

The fans of many turbofan engines have fan blades with a large slope atthe root of the fan airfoils and a large change in radius, from theleading edge to the trailing edge, at the fan blade roots. Theseproperties may provide certain aerodynamic advantages. However, when fanchord is minimized for engine length/weight reasons, or fan blade rootis thickened for structural reasons, the resulting high slope at the fanblade root can compromise the downstream fan root flow.

Consequently, the flow downstream of the fan blade root in such fans cancarry large circumferential wake and thick end wall boundary layers.This can cause high incidence angle of the flow at the downstreamstators, which may cause undesirable effects, such as the initiation ofpremature stall on the fan core stators downstream of the fan.

SUMMARY

There is accordingly provided a compressor for a gas turbine engine, thecompressor comprising: a rotor and a stator located downstream of therotor, the stator comprising a plurality of vanes, each of the vaneshaving an airfoil extending generally radially from a root proximate aninner hub of the stator to a radially outer tip, a span of the airfoildefined between the root and the tip, the airfoil having a leading edge,a trailing edge, and a chord extending between the leading edge and thetrailing edge to define a chord length, the airfoil having a pressureside surface and a suction side surface each extending on opposite sidesof the airfoil between the leading edge and the trailing edge, and afillet disposed at the root of the airfoil, the fillet extendingupstream from the leading edge of a remainder of the airfoil, the filletextending away from the pressure side surface a greater distance thanthe fillet extends away from the suction side surface.

There is also provided a turbofan engine comprising a fan and a casingdefining a bypass duct surrounding an engine core defining an annulargas passage, a fan stator disposed within the engine core downstream ofthe fan, the fan stator including a plurality of vanes circumferentiallyspaced-apart around a circumference of fan stator within the annular gaspassage, each of the vanes having an airfoil extending between a rootand a tip spaced apart by a span of the airfoil, the airfoil having aleading edge and a trailing edge spaced apart along a chord line by achord length of the airfoil, a pressure side surface and a suction sidesurface respectively extending on opposite sides of the airfoil betweenthe leading edge and the trailing edge, and a leading edge filletdisposed at the root of the airfoil on the pressure side surface.

There is further provided a gas turbine engine comprising: a compressorwith a rotor and a stator located downstream of the rotor, the statorhaving a plurality of vanes each having an airfoil extending span-wisebetween a root proximate an inner hub of the stator and a tip, theairfoil extending chord-wise between a leading edge and a trailing edge,a fillet disposed at the leading edge of the root of the airfoil andextending between a pressure side surface of the airfoil and theradially inner hub.

BRIEF DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-sectional view of a gas turbine engine;

FIG. 2 is a schematic side elevational view of a vane of a compressorstator of the gas turbine engine of FIG. 1;

FIG. 3 is a schematic cross-sectional view of the vane of the compressorstator of FIG. 2, taken through line 3-3 in FIG. 2;

FIG. 4 is a schematic cross-sectional view of the vane of the compressorstator of FIG. 2, taken through line 4-4 in FIG. 2;

FIG. 5A is a partial side perspective view of the vane of the compressorstator of FIG. 2; and

FIG. 5B is a partial front perspective view of the vane of thecompressor stator of FIG. 2.

DETAILED DESCRIPTION

FIG. 1 illustrates a gas turbine engine 10 of a type preferably providedfor use in subsonic flight, generally comprising in serial flowcommunication a compressor 13 for pressurizing the air, a combustor 14in which the compressed air is mixed with fuel and ignited forgenerating an annular stream of hot combustion gases, and a turbinesection 15 for extracting energy from the combustion gases.

The compressor 13 includes one or more axial compressor stages 16 withinthe core 22 of the gas turbine engine 10. Although the gas turbineengine 10 may be of different types (turboprop, turboshaft, turbofan,etc.), in the embodiment depicted in FIG. 1 the gas turbine engine 10 isa turbofan engine, and therefore includes a fan 12, upstream of theengine core 22, through which ambient air enters the engine. The airpropelled downstream from the fan 12 is split into a bypass duct 24 andthe engine core 22.

Each compressor stage 16 within the engine core 22 includes one or morerows of compressor stators 17 located immediately downstream of a row ofcompressor rotors 18. A fan core stator 17A, located downstream of thebladed rotor of the fan 12, is also disposed within the engine core 22,near an inlet thereof and upstream from the axial compressor stages 16.It is understood that the fan 12 forms part of the compressor 13 of theengine 10, and comprises a first stage, low pressure compressor.Accordingly, the fan core stator 17A and the compressor stators 17 ofthe downstream axial compressor stages 16 all constitute compressorstators, as defined herein, and may therefore have the compressor statorvanes 20 as will be defined hereinbelow.

Each of the compressor stators 17,17A is a non-rotating component thatguides the flow of pressurized air downstream from the compressor corerotors 18 and/or the fan 12. The compressor rotors 18 and the fan 12rotate about a longitudinal center axis 19 of the gas turbine engine 10to perform work on the air.

Each of the compressor stators 17,17A comprises a plurality of statorvanes 20. Each stator vane 20 is a stationary body that diffuses theairflow impinging thereon, thereby converting at least some of thekinetic energy of the incoming airflow into increased static pressure.The stator vanes 20 also redirect the airflow toward the next downstreamcompressor rotor 18 and/or to the combustor (in the case of themost-downstream compression stage 16).

With reference now to FIGS. 2-5B, the stator vanes 20 of the compressorstators 17,17A of the compressor 13 will now be described in furtherdetail. It is understood that only one, or any two or more, of thestages of the compressor 13 may have the stator vanes 20 as describedhereinbelow. For example, only the fan core stator 17A may be formedwith the stator vanes 20 as described herein. Alternately, only one ormore of the stator vanes 17 of the compressor 13 may have the statorvanes 20 as described herein. Alternately still, both the fan corestator 17A and the stator vanes 17 may all have the stator vanes 20 asdescribed herein. The terms “upstream” and “downstream” as used hereinare made with reference to the flow F of air through the compressorstators 17,17A and thus the airfoil F over each vane 20 thereof.

Referring first to FIGS. 2-3, each stator vane 20 (or simply “vane” 20)of the present disclosure includes an airfoil 21 shaped and sized toeffect the above-describe functionality. The airfoil 21 extendsgenerally radially between a root 26, disposed adjacent to a radiallyinner hub 27 of the compressor stator 17,17A, and a distal tip 28,disposed adjacent to an outer shroud 29 of the compressor stator 17,17A.The airfoils 21 define a span S of the vane 20 which extends between theroots 26 and the tips 28. The airfoil 21 therefore extends substantiallyin a radial direction (i.e. in a direction that generally extendsparallel to a radial line from the center axis 19 of the gas turbineengine 10), which extends between the root 26 and the tip 28, and whichdefines the radial span S of the airfoil 21. This may also be referredto as the height of the vanes 20.

The chord length C of the airfoil 21 is defined between a leading edge30 of the airfoil 21, and a trailing edge 32 of the airfoil 21. Theextent of the airfoil 21 along its chord is therefore defined betweenthe leading edge 30 and the trailing edge 32, and is referred to hereinas the chord length C. In the depicted embodiment, the chord length C isthe length of the chord line, which may be thought of as a straight lineconnecting the leading and trailing edges 30 and 32.

The airfoil 21 also includes a pressure side surface 34 and a suctionside surface 36, disposed on opposite sides of the airfoil and eachextending between the leading edge 30 and trailing edge 32.

Referring still to FIGS. 2-3, the vane 20 further includes a fillet 40disposed at the root 26 of the airfoil 21, proximate the leading edge30. Each of the airfoils 21 accordingly has a root 26 with anasymmetrical fillet 40 on a leading edge pressure side of the airfoil.The asymmetrical filet 40 is such that it is larger on the pressure sideof the airfoil than on the suction side of the airfoil. The term“larger” as used herein in this context is understood to mean a filetthat is any one or more of taller, longer, wider, greater radius,thicker, greater volume, greater mass, or the like, and/or anycombination of these characteristics. In the embodiment of FIGS. 2-3,for example, the filet 40 on the pressure side 34 of the airfoil 21protrudes outwardly from the pressure side surface of the airfoil agreater extent than does any filet (if one is present at all) located onthe suction side 36 of the airfoil, and is therefore said to be largerthan any filet on the suction side 36. (In this embodiment, however, thesuction side 36 is substantially free of any filet.)

The fillet 40 may extend between the airfoil 21 and the radially innervane platform 27, as seen in FIGS. 2 and 3. In the depicted embodiment,the fillet 40 is disposed exclusively on the pressure side 34 of theairfoil. While the fillet 40 may extend onto the suction side surface 36of the airfoil 21 in certain alternate embodiments, in all cases thefillet 40 is larger on the pressure side surface 34 of the airfoil 21than on the suction side surface 36, such a way that the fillet 40 isconfigured to provide a greater aerodynamic effect on the pressure sideof the airfoil than on the suction side thereof. The term “larger” asused in this context is understood to mean any one of wider, higher,longer and/or thicker (or other corresponding sized-based physicalproperties). In the depicted embodiment, the fillet 40 is disposed onlyon the pressure side 34 only of the airfoil 21 (i.e. the fillet 40 doesnot extend around to the suction side surface 36 of the airfoil 21, suchthat the suction side 36 is free of any fillet 40).

The fillet 40 is disposed at the leading edge 30 of the airfoil 21, andextends at least partially along the leading edge a span-wise distanceaway from the root 26 of the airfoil 21. In one particular embodiment,the fillet 40 extends radially away from the root 26 a span-wisedistance H of less than 10% of the total span S of the airfoil 21. Moreparticularly still, the span-wise distance H of the fillet 40 may befrom 2 to 10% of the total span S of the airfoil 21, beginning at theroot 26 thereof.

As best seen in FIGS. 2 and 4, the fillet 40 of the depicted embodimentaxially extends upstream and/or downstream relative to the leading edge30 of the airfoil 21. It is understood that the leading edge 30 as usedin this context (i.e. as a stream-wise reference point for the fillet40) corresponds to the leading edge 30 of a majority of the airfoil 21of the vane 20 outside the region of the vane having the fillet 40. Morespecifically, if a radially extending axis 50 is axially (i.e.stream-wise) aligned with the leading edge 30 of a major portion of theairfoil outside the fillet region, this defines the location of the“leading edge 30” as used as a reference point for the axially extendingextent of the fillet 40.

The fillet 40 may accordingly extend upstream and/or downstream relativeto the leading edge 30 of the airfoil 21, on the pressure side 34 of thevane 20.

As seen in the embodiment of FIGS. 2 and 4, the fillet 40 axiallyextends upstream a distance A from the leading edge 30 (and the axis50), and axially extends downstream a distance B from the leading edge30 (and the axis 50).

In one embodiment, the fillet 40 is disposed only on an upstream half ofthe pressure side surface 34, and therefore the distance B may be lessthan 50% of the total chord length C. In a more particular embodiment,the distance B may be from 10% to 50% of the total chord length C.

In one embodiment, the fillet 40 is disposed only within a radiallyinnermost portion of the airfoil 21, proximate the root 26, andtherefore the distance H may be less than 10% of the total span S. In amore particular embodiment, the distance H may be from 2% to 10% of thetotal span S.

In one embodiment, the fillet 40 extends upstream of the leading edge 30(and therefore the axis 50) a distance A that is less than 20% of thetotal chord length C. In a more particular embodiment, the distance Amay be from 5% to 20% of the total chord length C.

Accordingly, the fillet 40 in effect extends the leading edge of theairfoil 21, at its root 26, upstream (e.g. by 5-20% of the airfoil chordat the hub) relative to a remainder of the airfoil outside the filletregion. This has the effect of increasing the staggered angle at theroot 26 of the airfoil 21 of the vane 20. The staggered angle θ is theangle defined between a horizontal axis X and a camber line CL of theairfoil, at any given span-wise location on the leading edge 30.

As best seen in FIG. 4, the staggered angle θ_(F) at the fillet 40 isgreater than a staggered angle θ_(A) of a majority of the airfoil 21 atthe leading edge 30 outside of the fillet region. More specifically, thestaggered angle θ_(F) of the fillet 40 may be between 5 and 10 degreesgreater than the staggered angle θ_(A) of the airfoil 21 outside thefillet region. By increasing the staggered angle θ_(F) at the leadingedge of the root 26 of the vane airfoil 21, due to the presence of thefillet 40, the incidence angle of the airflow F is reduced (relative tothat over the remainder of the airfoil, outside the filleted region).

The fillet 40 may accordingly help improve aerodynamic performance andflow quality downstream of the stator 17,17A, and may increase the stallrange of the compressor 13. A reduction in stator flow separation mayalso result, which can lead to performance improvements for thedownstream rotor(s).

Referring still to FIG. 4, the fillet 40 may have a fillet radius R thatis from 5% and 20% of the total span S of the airfoil 21.

As also best seen in FIG. 4, in order to create the fillet 40, thesuction side surface 36 of the airfoil 21 is extended upstream (from theleading edge 30), to thereby form the higher staggered angle θ_(F). Thesuction side surface 36 is extended smoothly upstream along the suctionsurface by the distance A, which thereby increases the higher staggeredangle θ_(F) relative to the staggered angle θ_(A) of the airfoil 21outside the fillet region.

As can be seen in FIGS. 5A and 5B, the fillet 40 on the pressure side ofthe airfoil 21 may be shaped with a blended curve (i.e. a so-called“variable” curve fillet), so as to smoothly (in an aerodynamic sense)blend the leading edge extension formed by the fillet 40 smoothly backto/with the pressure side surface 34 and the leading edge 30 of theairfoil 21 and the inner hub of the stator. This variable fillet shapeon the pressure side may help to speed up the pressure side airflow, andreduce secondary flow losses. For example, this may reduce transverseflow, which results in less surface boundary layer radial flowmigration. This “variable” or blended fillet 40 accordingly helps tospeed up the pressure side flow and reduce secondary flow. The vanes 20having this fillet 40 may also cause less transverse flow, which canresult in less surface boundary layer radial flow migration. Overallstage performance of the compressor 13 may as a result be improved(lower stator losses and better aerodynamic performance).

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.Still other modifications which fall within the scope of the presentinvention will be apparent to those skilled in the art, in light of areview of this disclosure, and such modifications are intended to fallwithin the appended claims.

The invention claimed is:
 1. A compressor for a gas turbine engine, thecompressor comprising: a rotor and a stator located downstream of therotor, the stator comprising a plurality of vanes, each of the vaneshaving an airfoil extending generally radially from a root proximate aninner hub of the stator to a radially outer tip, a span of the airfoildefined between the root and the tip, the airfoil having a leading edge,a trailing edge, and a chord extending between the leading edge and thetrailing edge to define a chord length, the airfoil having a pressureside surface and a suction side surface each extending on opposite sidesof the airfoil between the leading edge and the trailing edge, and afillet disposed at the root of the airfoil, the fillet extendingupstream from the leading edge of a remainder of the airfoil, the filletextending away from the pressure side surface a greater distance thanthe fillet extends away from the suction side surface, a height of thefillet taken in a spanwise direction is maximum at the leading edge, thefillet having a width in a direction normal to the pressure side of theremainder of the airfoil, the height and the width of the filletdecreasing from the leading edge such that the fillet blends smoothlyinto the pressure side of the reminder of the airfoil in a downstreamdirection, the fillet being disposed only on the pressure side surfaceof the airfoil, the suction side surface being free of the fillet. 2.The compressor as defined in claim 1, wherein the fillet extendsdownstream from the leading edge.
 3. The compressor as defined in claim2, wherein the fillet extends downstream from the leading edge achord-wise distance of less than 50% of the chord length.
 4. Thecompressor as defined in claim 3, wherein the chord-wise distance isfrom 10% to 50% of the chord length.
 5. The compressor as defined inclaim 1, wherein the fillet extends upstream from the leading edge achord-wise distance of less than 20% of the chord length.
 6. Thecompressor as defined in claim 5, wherein the chord-wise distance isfrom 5% to 20% of the chord length.
 7. The compressor as defined inclaim 1, wherein the fillet extends radially from the root along theleading edge.
 8. The compressor as defined in claim 7, wherein thefillet extends radially away from the root a span-wise distance of lessthan 10% of the span of the airfoil.
 9. The compressor as defined inclaim 8, wherein the span-wise distance is from 2% to 10% of the span ofthe airfoil.
 10. The compressor as defined in claim 1, wherein thefillet has a fillet radius that is from 5% to 15% of the span of theairfoil.
 11. The compressor as defined in claim 1, wherein the filletblends smoothly with the pressure side surface and the leading edge ofthe airfoil, and the inner hub of the stator.
 12. The compressor asdefined in claim 1, wherein a first staggered angle is defined at thefillet, and a second staggered angle is defined at the leading edge ofthe airfoil at a point thereon outside of the fillet, the firststaggered angle being greater than the second staggered angle, andwherein the first staggered angle is from 5 to 10 degrees greater thanthe second staggered angle.
 13. A turbofan engine comprising a fan and acasing defining a bypass duct surrounding an engine core defining anannular gas passage, a fan stator disposed within the engine coredownstream of the fan, the fan stator including a plurality of vanescircumferentially spaced-apart around a circumference of the fan statorwithin the annular gas passage, each of the vanes having an airfoilextending between a root and a tip spaced apart by a span of theairfoil, the airfoil having a leading edge and a trailing edge spacedapart along a chord line by a chord length of the airfoil, a pressureside surface and a suction side surface respectively extending onopposite sides of the airfoil between the leading edge and the trailingedge, and a leading edge fillet disposed at the root of the airfoil onthe pressure side surface, the leading edge fillet having a height takenin a spanwise direction and a width in a direction normal to thepressure side, the height and the width decreasing from the leading edgesuch that the fillet blends smoothly into the pressure side of thereminder of the airfoil in a downstream direction, a first staggeredangle defined at the fillet and a second staggered angle is defined atthe leading edge of the airfoil at a point thereon radially above thefillet, the first staggered angle being from 5 to 10 degrees greaterthan the second staggered angle.
 14. The turbofan engine of claim 13,wherein the fillet extends upstream from the leading edge by a distanceof 5-20% of the chord length.
 15. The turbofan engine of claim 13,wherein the fillet extends downstream from the leading edge by adistance of 10-50% of the chord length.
 16. The turbofan engine of claim13, wherein the fillet extends radially away from the root by a distanceof 2-10% of the span.
 17. A compressor for a gas turbine engine, thecompressor comprising: a rotor and a stator located downstream of therotor, the stator comprising a plurality of vanes, each of the vaneshaving an airfoil extending generally radially from a root proximate aninner hub of the stator to a radially outer tip, a span of the airfoildefined between the root and the tip, the airfoil having a leading edge,a trailing edge, and a chord extending between the leading edge and thetrailing edge to define a chord length, the airfoil having a pressureside surface and a suction side surface each extending on opposite sidesof the airfoil between the leading edge and the trailing edge, and afillet disposed at the root of the airfoil, the fillet extendingupstream from the leading edge of a remainder of the airfoil, the filletextending away from the pressure side surface a greater distance thanthe fillet extends away from the suction side surface, a height of thefillet taken in a spanwise direction is maximum at the leading edge, thefillet having a width in a direction normal to the pressure side of theremainder of the airfoil, the height and the width of the filletdecreasing from the leading edge such that the fillet blends smoothlyinto the pressure side of the reminder of the airfoil in a downstreamdirection, wherein a first staggered angle is defined at the fillet, anda second staggered angle is defined at the leading edge of the airfoilat a point thereon outside of the fillet, the first staggered anglebeing greater than the second staggered angle, and wherein the firststaggered angle is from 5 to 10 degrees greater than the secondstaggered angle.